Gas turbine engine oil buffering

ABSTRACT

A turbine engine includes a shaft, a fan, at least one bearing mounted on the shaft and rotationally supporting the fan, a fan drive gear system coupled to drive the fan, a bearing compartment around the at least one bearing and a source of pressurized air in communication with a region outside of the bearing compartment.

CROSS REFERENCE TO RELATED APPLICATIONS

The present disclosure is a continuation in part of U.S. patentapplication Ser. No. 12/708,621, filed Feb. 19, 2010.

BACKGROUND OF THE INVENTION

The present invention relates generally to gas turbine engines, and moreparticularly, to a system for supplying buffer air and ventilation airto the bearing compartments and shaft(s) in gas turbine engines.

In typical multi-shaft bypass jet turbine engines, one shaft supportsthe rotors of a low pressure compressor and a low pressure turbine andanother shaft supports the rotors of a high pressure compressor and ahigh pressure turbine. Generally, each of the shafts is supported bybearings, and each bearing is lubricated by a forced lubrication systemwhich circulates lubricating oil fed by a pump.

In the forced lubricating system, high pressure air is drawn from thehigh pressure compressor and is conducted to the exterior of the oilseals of the bearing compartments to keep the interior of the bearingcompartments at a lower pressure than its immediate surroundings. Thispressure differential prevents the lubricating oil from leaking out ofthe bearing compartments. In particular, high pressure buffer air drawnfrom the high pressure compressor is utilized because at least one ofthe bearing compartments is located in a high pressure environment wherebuffer air from the low pressure compressor would not provide adequatecompartment pressurization at low power engine operating conditions.Unfortunately, buffer air drawn from the high pressure compressor isexcessively hot and requires cooling at higher power engine operatingconditions. Therefore, a dedicated cooler is required to lower thetemperature of the buffer air. This cooler adds additional weight to theengine and can be difficult to package especially in smaller enginemodels.

SUMMARY OF THE INVENTION

A disclosed example turbine engine according to an exemplary embodimentincludes a shaft, a fan, at least one bearing mounted on the shaft androtationally supporting the fan, a fan drive gear system coupled todrive the fan, a bearing compartment around the at least one bearing,and a source of pressurized air in communication with a region outsideof the bearing compartment.

In a further embodiment of the foregoing turbine engine, the fan drivegear system includes an epicyclic gear train.

In a further embodiment of the foregoing turbine engine, the epicyclicgear train has a gear reduction ratio of greater than or equal to about2.3.

In a further embodiment of the foregoing turbine engine, the turbineengine, the epicyclic gear train has a gear reduction ratio of greaterthan or equal to 2.3.

In a further embodiment of the foregoing turbine engine the epicyclicgear train has a gear reduction ratio of greater than or equal to about2.5.

In a further embodiment of the foregoing turbine engine, the epicyclicgear train has a gear reduction ratio of greater than or equal to 2.5.

In a further embodiment of the foregoing turbine engine, the fan definesa bypass ratio of greater than about ten (10) with regard to a bypassairflow and a core airflow.

In a further embodiment of the foregoing turbine engine, the fan definesa bypass ratio of greater than 10.5:1 with regard to a bypass airflowand a core airflow.

In a further embodiment of the foregoing turbine engine, the fan definesa bypass ratio of greater than ten (10) with regard to a bypass airflowand a core airflow.

In a further embodiment of the foregoing turbine engine, the fan definesa pressure ratio that is less than about 1.45.

In a further embodiment of the foregoing turbine engine, the fan definesa pressure ratio that is less than 1.45.

In a further embodiment of the foregoing turbine engine, a corecompressor section is the source of the pressurized air.

A disclosed method of operating a turbine engine that includes a shaft,a fan, at least one bearing mounted on the shaft and rotationallysupporting the fan, a fan drive gear system coupled to drive the fan anda bearing compartment around the at least one bearing according to anexemplary embodiment includes the steps of providing pressurized air toa region outside of the bearing compartment to establish a positivepressure differential between the region outside of the bearingcompartment and the interior of the bearing compartment.

In a further embodiment of the foregoing method the fan drive gearsystem includes an epicyclic gear train.

In a further embodiment of the foregoing method, the epicyclic geartrain has a gear reduction ratio of greater than or equal to about 2.3.

In a further embodiment of the foregoing method, including the step ofproviding the pressurized air from a core compressor section.

Although different examples have the specific components shown in theillustrations, embodiments of this invention are not limited to thoseparticular combinations. It is possible to use some of the components orfeatures from one of the examples in combination with features orcomponents of another of the examples.

These and other features disclosed herein can be best understood fromthe following specification and drawings, the following of which is abrief description.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic partial sectional view of a gas turbine enginewith a centrifugal compressor driven by an accessory gearbox.

FIG. 2 is a schematic of the buffer and ventilation air system of thegas turbine engine of FIG. 1.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT

The present application describes new pressurization and ventilationsystems for bearing compartments and shafts of a gas turbine engine. Inparticular, the present application describes an assembly and a methodfor providing buffer and/or ventilation air to bearing compartmentsand/or shafts of a gas turbine engine. The gas turbine engine describedincludes a dedicated centrifugal compressor that compresses bleed airfrom a low pressure compressor section and/or a fan section of the gasturbine engine. The compressed air is delivered to the core of theengine to buffer the bearing compartments and/or ventilate one or moreshafts. The centrifugal compressor compresses the bleed air to a higherpressure more efficiently than traditional axial compressor arraysbecause it avoids the loss of kinetic energy and the throttling lossesat the compressor case that are experienced with traditional axialcompressor arrays. By utilizing the centrifugal compressor, air is drawnfrom the fan section and/or the low pressure compressor section andcompressed to a desired optimal pressure and temperature, therebyeliminating the need for cooling within a cooler. This arrangement alsoreduces the likelihood of an inadequately pressurized bearingcompartment at low power engine operating conditions, discussedpreviously. The flow rate and pressure ratio requirements of thecentrifugal compressor are low enough to allow for a compact design thatcan fit within various locations such as the engine core and allow thecentrifugal compressor to be integrated as an accessory to be driven bythe gearbox.

FIG. 1 shows a schematic partial cross section of a portion of a gasturbine engine 10. Gas turbine engine 10 has bearing compartments 12Aand 12B which house anti-friction bearings that support shafts 14A and14B. Gas turbine engine 10 is defined around an engine centerline CLabout which various engine sections rotate. In FIG. 1, only a portion ofgas turbine engine 10 including a rotor section 16, a fan section 18, alow pressure compressor (LPC) section 20, and a high pressure compressor(HPC) section 22 is illustrated. Gas turbine engine 10 is illustrated asa high bypass ratio turbofan engine with a dual spool arrangement inwhich fan section 18 and LPC 20 are connected to a low pressure turbinesection (not shown) by rotor 16, fan drive gear system 15, and shaft14A, and high pressure compressor section 22 is connected to a highpressure turbine section (not shown) by second shaft 14B. The generalconstruction and operation of gas turbine engines, and in particularturbofan engines, is well-known in the art, and therefore, detaileddiscussion herein is unnecessary. It should be noted, however, thatengine 10 is shown in FIG. 1 merely by way of example and notlimitation. The present invention is also applicable to a variety ofother gas turbine engine configurations, such as a turbofan enginewithout fan-drive gear system and a turboprop engine, for example.

In one example, the gas turbine engine 10 is a high-bypass gearedarchitecture aircraft engine. In one disclosed, non-limiting embodiment,the engine 10 has a bypass ratio that is greater than about six (6) toten (10), the fan drive gear system 15 is epicyclic gear train andincludes a planetary gear system or other gear system with a gearreduction ratio of greater than about 2.3 or greater than about 2.5, anda low pressure turbine of the engine 10 has a pressure ratio that isgreater than about 5. In one disclosed embodiment, the engine 10 bypassratio is greater than about ten (10:1) or greater than about 10.5:1, thea fan rotor 24 diameter is significantly larger than that of the lowpressure compressor of the compressor section 20/22, and the lowpressure turbine has a pressure ratio that is greater than about 5:1. Inone example, the epicyclic gear train has a gear reduction ratio ofgreater than about 2.3:1 or greater than about 2.5:1. It should beunderstood, however, that the above parameters are only exemplary of oneembodiment of a geared architecture engine and that the presentinvention is applicable to other gas turbine engines including directdrive turbofans.

A significant amount of thrust is provided by a bypass flow B due to thehigh bypass ratio. The fan of the engine 10 is designed for a particularflight condition—typically cruise at about 0.8M and about 35,000 feet.The flight condition of 0.8 M and 35,000 ft, with the engine at its bestfuel consumption—also known as “bucket cruise TSFC”—is the industrystandard parameter of 1 bm of fuel being burned divided by 1 bf ofthrust the engine produces at that minimum point. “Low fan pressureratio” is the pressure ratio across the fan blade alone. The low fanpressure ratio as disclosed herein according to one non-limitingembodiment is less than about 1.45. “Low corrected fan tip speed” is theactual fan tip speed in ft/sec divided by an industry standardtemperature correction of [(Tambient deg R)/518.7)̂0.5]. The “Lowcorrected fan tip speed” as disclosed herein according to onenon-limiting embodiment is less than about 1150 ft/second.

In addition to fan section 18, low pressure compressor section 20, andhigh pressure compressor section 22, gas turbine engine 10 includes afan rotor 24, an outer case 26, a bypass duct 27, an inner fan case 28,an intermediate case 30, an accessory gearbox 32, a centrifugalcompressor 34, and a radial drive shaft 36. Inner fan duct 28 andintermediate case 30 define a core compartment 38.

Air enters fan section 18 of turbine engine 10 where it is divided intostreams of primary air Ar and bypass air AB after passing through thefan 24. Bypass air AB flows inside outer case 26 and along inner fancase 28 and eventually exits bypass duct 27 and engine 10. The fan 24 isdisposed radially inward of outer case 26 and is rotated by the lowpressure turbine (not shown) through the shaft 14 and fan-drive gearsystem 15 to accelerate the bypass air AB through fan section 18,thereby producing a significant portion of the thrust output of engine10.

The primary air Ar (also known as gas path air) is directed firstthrough the low pressure compressor section 20 (which is partiallybounded by intermediate case 30) and then through high pressurecompressor section 22. As illustrated in FIG. 1, accessory gearbox 32 isconnected to intermediate case 30 and extends radially outward of enginecenterline CL away from low pressure compressor section 20. As is knownin the art, the location of accessory gearbox 32 in FIG. 1 is by way ofexample and not limitation. In other embodiments, accessory gearbox 32can be disposed on outer case 26, inner fan case 28 or in otherlocations including within the core of the gas turbine engine 10.Accessory gearbox 32 is connected to and drives centrifugal compressor34. More particularly, accessory gearbox 32 transfers torque from radialdrive shaft 36 to centrifugal compressor 34. Radial drive shaft 36 iscoupled to accessory gearbox 32 and extends into bearing compartment 12B(specifically called a high rotor thrust bearing compartment) to couplewith and transfer torque from shaft 14B.

Accessory gearbox 32 commonly drives various engine accessoriesincluding an electrical generator (not shown) and a main engine oilsystem, which is used to lubricate components of the engine includingthe bearings. Accessory gearbox 32 is specifically adapted to drivecentrifugal compressor 34, which is disposed within core compartment 38.Scoops or other known devices bleed air at a lower pressure from fansection 18 and/or low pressure compressor section 20. This bleed air isdirected to centrifugal compressor 34 where it is compressed to a higherpressure. Thus, lower pressure bleed air directed to the centrifugalcompressor 34 can comprise either bypass air AB or primary air Ar or amixture of both. Primary air Ar can be drawn off various stages of thelow pressure compressor section 20 as desired.

Centrifugal compressor 34 operates in a manner known in the art tocompress lower pressure bleed air to a higher pressure. Operation of thecentrifugal compressor causes a pressure differential that circulatesthe higher pressure air A to all bearing compartments including forwardbearing compartments 12A and 12B and more rearward bearing compartments(FIG. 2) to act as buffer air. In addition to or in alternative to beingused as buffer air for the bearing compartments, higher pressure air Acan be used as ventilation air to ventilate shaft 14 in a manner know inthe art.

A pressure differential between the interior of bearing compartments 12Aand 12B and higher pressure air A, along with the configuration of thebearing compartment seals, allows higher pressure air A to migrateacross bearing compartment seals into the bearing compartments 12A and12B. Migration of higher pressure air A across seals helps to preventthe caustic and flammable lubricating oil from leaking out of bearingcompartments 12A and 12B. Utilization of centrifugal compressor 34allows higher pressure A to be compressed to a desired optimal pressureand temperature to provide adequate pressurization to bearingcompartments 12A and 12B to prevent oil leakage therefrom. Compressingair A to the desired optimal temperature and pressure eliminates theneed for the cooling of air A within a cooler, thereby reducing engine10 weight and providing more design space within engine 10.

FIG. 2 shows a schematic view of a system 39 that provides buffer airand ventilation air within gas turbine engine 10. System 39 includes alower pressure location 40, lower pressure external lines or internalpassages 42, higher pressure external lines or internal passages 44,metering devices 46, structures 48A-48E, bearing compartments 12A-12Dsurrounded by seal cavities 13A-13F, and a secondary air flow 50.

Accessory gearbox 32 is coupled to and acts to drive centrifugalcompressor 34. Scoops or other known devices bleed air at a lowerpressure from lower pressure location 40 within gas turbine engine 10.In one embodiment, location 40 comprises fan section 18 (FIG. 1) and/orlow pressure compressor section 20 (FIG. 1). Bleed air is directed asair flow through lower pressure external lines or internal passages 42to centrifugal compressor 34, which operates to compress the lowerpressure bleed air to a higher pressure. The higher pressure air that iscompressed in centrifugal compressor 34 circulates away from centrifugalcompressor 34 through higher pressure external lines and internalpassages 44. Higher pressure air flow branches into several flows thatare directed through one or more metering devices 46, such as valves ororifices, which throttle air flow as desired. In the embodimentillustrated in FIG. 2, higher pressure air flow is directed throughstructures 48A-48E. Structures 48A-48E can variously comprise strutsand/or other portions of the front center body, intermediate case, ormid-turbine frame of gas turbine engine 10. Higher pressure air flowpasses through structures 48A-48E to provide buffer air to seal cavities13A-13F that surround bearing compartments 12A-12D and ventilation airto shaft 14 (FIG. 1). In the embodiment shown, secondary air flow 50continues from seal cavity 13C as ventilation air along the innerdiameter of shaft 14A. Secondary air flow 50 also provides buffer air toseal cavity 13F that buffers bearing compartment 12D toward the rear ofgas turbine engine 10.

While the invention has been described with reference to an exemplaryembodiment(s), it will be understood by those skilled in the art thatvarious changes may be made and equivalents may be substituted forelements thereof without departing from the scope of the invention. Inaddition, many modifications may be made to adapt a particular situationor material to the teachings of the invention without departing from theessential scope thereof. Therefore, it is intended that the inventionnot be limited to the particular embodiment(s) disclosed, but that theinvention will include all embodiments falling within the scope of theappended claims.

What is claimed is:
 1. A turbine engine comprising: a shaft; a fan; atleast one bearing mounted on the shaft and rotationally supporting thefan; a fan drive gear system coupled to drive the fan; a bearingcompartment around the at least one bearing; and a source of pressurizedair in communication with a region outside of the bearing compartment.2. The turbine engine as recited in claim 1, wherein the fan drive gearsystem includes an epicyclic gear train.
 3. The turbine engine asrecited in claim 2, wherein the epicyclic gear train has a gearreduction ratio of greater than or equal to about 2.3.
 4. The turbineengine as recited in claim 2, wherein the epicyclic gear train has agear reduction ratio of greater than or equal to 2.3.
 5. The turbineengine as recited in claim 2, wherein the epicyclic gear train has agear reduction ratio of greater than or equal to about 2.5.
 6. Theturbine engine as recited in claim 2, wherein the epicyclic gear trainhas a gear reduction ratio of greater than or equal to 2.5.
 7. Theturbine engine as recited in claim 1, wherein the fan defines a bypassratio of greater than about ten (10) with regard to a bypass airflow anda core airflow.
 8. The turbine engine as recited in claim 1, wherein thefan defines a bypass ratio of greater than 10.5:1 with regard to abypass airflow and a core airflow.
 9. The turbine engine as recited inclaim 1, wherein the fan defines a bypass ratio of greater than ten (10)with regard to a bypass airflow and a core airflow.
 10. The turbineengine as recited in claim 1, wherein the fan defines a pressure ratiothat is less than about 1.45.
 11. The turbine engine as recited in claim1, wherein the fan defines a pressure ratio that is less than 1.45. 12.The turbine engine as recited in claim 1, including a core compressorsection, and the core compressor section is the source of thepressurized air.
 13. A method of operating a turbine engine, the methodcomprising: in a gas turbine engine that includes a shaft, a fan, atleast one bearing mounted on the shaft and rotationally supporting thefan, a fan drive gear system coupled to drive the fan and a bearingcompartment around the at least one bearing, providing pressurized airto a region outside of the bearing compartment to establish a positivepressure differential between the region outside of the bearingcompartment and the interior of the bearing compartment.
 14. The methodas recited in claim 13, wherein the fan drive gear system includes anepicyclic gear train
 15. The method as recited in claim 14, wherein theepicyclic gear train has a gear reduction ratio of greater than or equalto about 2.3.
 16. The method as recited in claim 13, including providingthe pressurized air from a core compressor section.